1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically for an air cooled turbine rotor blade with trailing edge cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a combustor that produces a hot gas stream and a turbine that reacts with the hot gas stream to produce mechanical work. The efficiency of the engine can be increased by passing a higher temperature gas into the turbine, referred to as the turbine inlet temperature. The turbine inlet temperature is limited to the material properties of the turbine, especially the first stage rotor blades and guide vanes, as well as to the amount of cooling for these airfoils. complex airfoil cooling circuit s have been proposed to provide for ever more increases in cooling capability while minimizing the amount of cooling air used to improve performance as well as increase part life.
Turbine blades and vanes are manufactured using the investment casting process in which a ceramic core representing the internal cooling passages is placed within a mold and liquid molten metal is poured into the mold. The mold includes a space in which the molten metal will flow and harden to represent the metallic portion of the airfoil. After the molten metal has solidified, the ceramic core is leached away, leaving the internal cooling air passages formed within the solidified metal. Additional machining can be required, for example to form the rows of film cooling holes that open onto the external surface of the airfoil.
FIGS. 1-7 show a prior art turbine rotor blade for an industrial gas turbine (IGT) engine with a leading edge region cooling circuit, a mid-chord region cooling circuit, and a trailing edge region cooling circuit. FIG. 2 shows a cross section side view of this circuit in which the mid-chord region cooling circuit includes a forward flowing three-pass serpentine flow circuit with a first leg 11 positioned adjacent to the trailing edge region cooling circuit to supply cooling air for it. The T/E region cooling circuit includes three rows of metering holes (21,22,23) that are staggered in the airfoil radial direction so as to produce a series of impingement cooling against the downstream rib, followed by a row of T/E exit holes or slots 24 to discharge the spent cooling air from the airfoil. FIG. 3 shows a diagram view of this cooling air circuit. FIG. 4 shows a close-up view of the T/E region cooling circuit with the first leg 11 of the forward flowing serpentine flow cooling circuit. FIG. 5 shows a cross section back view of a section through the first row of metering holes 21 as represented by line A-A in FIG. 4. As seen in FIG. 5, the row of metering holes extends from a lower continuous cooling channel 14 to an upper continuous cooling channel 13. The row of metering and impingement holes 21-23 (each hole is both a metering hole and an impingement hole) are of the same diameter as seen in FIG. 5.
The leading edge flow circuit provides cooling primarily for the leading edge which is the critical part of the blade from a durability spent point. Cooling air is fed into the airfoil through a single pass radial channel. Skewed trip strips are used on the pressure and suction inner walls of the radial cooling channel to augment the internal heat transfer performance. A multiplicity of impingement jets from the cooling supply channel pass through a row of cross-over metering holes in a first partition rib to provide backside impingement cooling for the blade leading edge inner surface. These cross-over holes are designed to support the leading edge ceramic core during casting of the blade, including removal of ceramic core material during a leaching process. The spent impingement cooling air is then discharged through a series of small diameter showerhead film cooling holes at a relative radial angle with the leading edge surface. A portion of the impingement air is also discharged through rows of pressure side and suction side gill holes. Therefore, a combination of impingement, convection and film cooling produces a blade leading edge metal temperature within acceptable levels. The castability of this arrangement has been demonstrated. In addition, multiple compartments can also be used in the leading edge impingement channel to regulate the pressure ration across the leading edge showerhead, eliminating showerhead film blow-off problems, and achieving optimum cooling performance with adequate backflow pressure margin and minimum cooling flow.
One major problem with air cooled turbine airfoils such as that in FIGS. 1-7 is that the ceramic core, which is made of a very brittle ceramic material, can shift during the casting process or even break. When the relatively heavy molten metal is poured into the mold and flows around the ceramic core, the heavy molten metal can shift the core into a position that will produce a defective casting. Or, some of the very fine ceramic pieces within the core can even break in half, resulting in what should be a cooling air passage to become a blocked passage. This is the main problem with the very fine cooling passages such as those formed as the metering holes in the T/E region cooling circuit.
Applicant has discovered that the temperature profile for the T/E cooling circuit varies from the root to the blade tip. FIG. 7 shows a blade relative gas temperature profile for the blade of FIG. 1 where FIG. 6 shows the T/E region cooling circuit and FIG. 7 shows a graph of the temperature versus the blade span height with the blade tip on the top and the blade root on the bottom. What is important in the FIG. 7 graph is that the peak temperature occurs around the middle portion of the blade span height. Thus, additional cooling is required for the airfoil mean section to achieve a proper sectional or local metal temperature.